The invention relates to a method of providing protection against the effects of solar originating protons, and to an attitude control system for a satellite implementing the method.
It is necessary to know the attitude of a satellite in order to control maneuvers relating to the mission the satellite is to perform.
In known manner, the attitude of a satellite can be determined by means of a relationship between the three axes of the satellite and reference points such as celestial stars, and this can be implemented by means of a star sensor installed onboard the satellite.
The present invention relates more particularly to so-called xe2x80x9c3-axisxe2x80x9d sensors that are self-contained.
In order to perform a 3-axis measurement, such sensors need to process information from at least two stars. In practice, they process a larger number, typically five to ten. The principle of detecting proton flux is based on the fact that there are at least N stars used for determining attitude.
Star sensors conventionally use an array or matrix of charge-coupled device (CCD) detectors. They also comprise a lens system or xe2x80x9ctelescopexe2x80x9d enabling an image to be obtained of a zone of space containing the star. Redundant equipment is generally provided to mitigate failures of the nominal equipment.
The field of view of a star sensor is determined by the dimensions of the array in the focal plane of the lens system. A system of coordinates can be given to the field of view or to the surface of the array of detectors, and the origin of this system of coordinates is preferably in the middle of the field of view. By knowing the positions of the stars seen relative to reference coordinates for the window of the sensor, and by correlating the resulting star xe2x80x9cpatternxe2x80x9d with a catalog of patterns, it is possible to obtain the attitude of the sensor, and thus of the satellite.
For better understanding of the invention, there follows an outline of the operation of a 3-axis sensor:
the sensor possesses one or more onboard star catalogs;
the sensor is characterized by two main modes of operation:
a) Acquisition mode: this mode enables the attitude of the sensor relative to an inertial frame of reference to be determined without a priori knowledge of said attitude. This initial determination is obtained by:
reading the entire CCD matrix;
determining the positions of candidate stars;
calculating the angular distances between pairs of candidate stars;
identifying candidate stars by comparing the calculated angular distances with an onboard catalog of pairs; and
determining the 3-axis attitude as a function of the coordinates of the stars that have been identified.
b) Tracking mode: this mode makes it possible to calculate 3-axis attitude finely as a function of measuring the positions of stars in the field of view. This mode relies on a priori knowledge of attitude with accuracy of the order of 0.3xc2x0 (depending on the sensors). This a priori knowledge is obtained by extrapolation from the preceding measurement. It is also possible to provide it by telemetry. It is this telemetry feature which is made use of below.
In tracking mode, when attitude is known a priori, it is possible to predict the positions of the stars that are being used and thus to read only those pixels which are adjacent thereto.
During a solar eruption, a proton flux is emitted that exceeds normal proton flux by several orders of magnitude. The interaction of such protons with the CCD matrix of a star sensor gives rise to electrons being deposited, and that can be interpreted as a star signal.
When the star sensor is in tracking mode, this phenomenon is not very critical since only a window of small size around each star is used in the processing, thereby minimizing the risk of interaction.
In contrast, when implementing acquisition mode, the entire CCD matrix is processed, and the presence of false stars runs the risk of disturbing or saturating the algorithm for recognizing the pattern of stars present.
There is then a risk of the failure detection isolation and recovery (FDIR) logic of the satellite triggering reconfiguration of the star sensor during a solar eruption, and that might make it impossible to reconfigure a redundant star sensor.
In this context, it is recalled that sensor reconfiguration can be due to two origins:
either the FDIR logic detects malfunction on the nominal sensor (by processing its telemetry), considers it to be broken down, and orders reconfiguration on the redundant equipment;
or else the FDIR logic detects a more general problem (e.g. excessive attitude error) but for which it is not possible to isolate the cause simply, in which case it orders general reconfiguration of all of the equipment in operation at that time, including the star sensor, even if it is functioning properly.
The resulting interruption of the main mission of the satellite will then be of a duration that corresponds to the typical duration of a solar eruption which is 48 hours (h) to 72 h. Such an interruption represents unacceptable lack of security and high extra cost.
The present invention enables this problem to be solved. It makes it possible to avoid interrupting a mission even if the active star sensor stops during an eruption of solar origin protons.
More particularly, the present invention provides a method of providing protection against the effects of solar originating protons, so as to make it possible for a star sensor to be reconfigured, said sensor being suitable for providing attitude data of a satellite, the method comprising the following steps:
detecting the arrival of an eruption of solar origin protons; and
triggering tracking mode operation of a second star sensor when the arrival of an eruption of solar origin protons has been detected so as to have at least one sensor in operation throughout the duration of the solar eruption.
According to another characteristic, tracking mode operation of the second sensor is obtained on the basis of attitude data supplied by the first sensor.
Detection comprises the following steps:
recording the number of stars rejected as obtained by processing the telemetry data of the first sensor;
determining a mean value for said number over a determined duration in order to eliminate false alarms;
comparing said mean value with a predetermined threshold; and
triggering operation of the second sensor when said threshold is reached or exceeded.
The invention also provides a system for controlling the attitude of a satellite having at least two star sensors onboard the satellite, a first of which is in operation, the system being characterized in that it comprises:
means for detecting the arrival of an eruption of solar origin protons, said means delivering a control signal; and
means for triggering operation of the second star sensor in tracking mode based on the attitude data provided by the first star sensor, said means being activated by said control signal.
According to another characteristic, the means for detecting the arrival of an eruption of protons of solar origin comprise means for recording the number of rejected stars as obtained by processing telemetry data from the star sensor in operation.
The means for detecting the arrival of an eruption of solar origin protons includes means for eliminating false alarms. Typically, these means comprise a lowpass filter.
According to anther characteristic, the means for detecting the arrival of an eruption of solar origin protons comprise decision logic having a predetermined maximum threshold beyond which it causes operation of the second star sensor to be triggered.